The present invention refers in general to a tooling for the manufacture of composite (CFRP) structures for aircraft, such as stringers, torsion boxes, skin panels, wing surfaces, horizontal tail or vertical stabilizers (HTP & VTP), etc.
More in particular, it is an object of the present invention to provide a tooling capable of satisfactorily curing pieces of CFRP, having parts with minimum thickness and/or very aggressive thickness changes.
The use of composite materials formed by an organic matrix and unidirectionally orientated fibers, such as Carbon Fiber Reinforced Plastic (CFRP), to construct several structural components of an aircraft, for example fuselage skin panels, torsion boxes, stringers, ribs, spars etc., is well known in the aeronautical industry.
Typically, skin panels are stiffened with a plurality of stringers (stiffeners) longitudinally arranged, so that the stringers improve the strength and buckling behavior of the skin panels. The stringers are conventionally bonded to the skin panel by co-curing or by co-bonding both parts together, or simply by applying a layer of adhesive (secondary-bonding).
The end of the stringer causes a redistribution of the loads being supported by the stringer and the skin. This produces two effects:                while the bending of the reinforced cover produces traction and compression loads, this punctual change in the structure of the stringer (stringer run-out), causes a moment in its termination point which tends to de-bond the bonding line between stiffener and skin.        at the same time, the redistribution of the loads has to be achieved through the bonding line in order to transfer it from the stringer to the skin at the stringer run-out area. At high load levels (as experimented for example in the wing cover) the resistance of the bonding lines could be compromised.        
The co-bonded union between skin and stringers in the torsion boxes of aircraft wings has to withstand loads in the magnitude of tons, which are at the limit of the structural capacity of said union at some critical spots, like the stringer run out area. These co-bonded unions can crack at this spot at high loads causing peeling loads due to two effects: elimination of the stringer web which causes a high load peak; and the main stringer termination, which produces a peak in the shear loads. In a typical configuration, these scenarios occur at the same time, penalizing the structural behavior of the bonding line.
To overcome these problems associated with the load transfer at stringer run-outs, it is well-known in this technical field, to reduce the total cross-sectional area of the stringer, usually by means of reducing the height of the stringer web (typically by cutting off a piece of it), and by progressively reducing the thickness of the stringer foot towards the run-out, by sequentially reducing the number of plies (dropping plies) towards the run-out.
FIG. 1 shows a conventional run-out section of a stringer (1) joined to a skin panel (2), wherein the stringer (1) has a T-shaped cross-section formed by a web (3) and a foot (4). As it can be observed in drawings (a, b) of FIG. 1, the thickness of the web (3) and foot (4) is reduced progressively towards the end of the stringer in order to reduce the cross-sectional area of the same, thereby reducing the elastic module of this run-out section.
Examples of this conventional solution are described in more detail for example in US patent applications US 2005/0211846 A1 and US 2012/0100343 A1.
However, existing metallic tools for curing pieces of composite material are not capable of satisfactorily manufacturing stringers with abrupt thickness reductions, as required for example for the web and foot of a run-out section. The existing metallic tooling is only capable of manufacturing ply drop offs of typically 1:200 (slope), which is not enough for some demanding applications.
If a metallic tool were to be used to cure a part with a ply drop-off bigger than 1:200, porosity problems would appear in the resulting piece, affecting the quality of the same. Tooling manufacturing tolerances and positioning tolerances due to slippage between the tool and the composite ply stack (5) would create gaps and misalignments between the tool fixed in place and the ply stacks (5) to be cured to form the composite material stringer.
FIG. 2 illustrates how these porosity problems are originated by gaps (7) created between the fresh stack of plies (5) of composite material, and hard tooling areas of the metallic tool (6) in contact with said plies (5) once the metallic tool is positioned for the curing process. During this curing process, a fresh (not-cured) formed stack of plies (5) of carbon fiber is heated and pressed against a metallic mold (5) by the application of a vacuum, so that the stack of plies (5) is cured and at the same time compacted by the metallic tool (6).
However, if there is a gap (7) between the working surface (8) (surface in contact with the stack of plies (5)) of the metallic tool (6) and the stack of plies (5), the uncured thermosetting resin of the composite material of the plies (5) flows towards those gaps (7) due to the effect of the applied vacuum. If the gap is too big, the resin cannot fill the entire gap originating said porosity problems in the resulting piece. This gap (7) is considered too big when the slope of the tooling is bigger than 1:200 for the typical manufacturing tolerances in a thickness of the typical stringers.
The same limitations affect the manufacturing of other CFRP structural elements with abrupt thickness change.